Electro-pneumatic environmental control system air circuit

ABSTRACT

An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

CROSS-REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Patent ApplicationNo. 62/432110 filed Dec. 9, 2016, and U.S. Non-Provisional PatentApplication Ser. No. 15/809244, filed on Nov. 10, 2017.

TECHNICAL FIELD

The present disclosure relates generally to aircraft air systems, andmore specifically to an air circuit for providing air to anenvironmental control system.

BACKGROUND

Aircraft, such as commercial airliners, typically include multiple gasturbine engines configured to generate thrust. The gas turbine enginesinclude a compressor section that compresses air, a combustor sectionthat mixes the air with a fuel and ignites the mixture, and a turbinesection across which the resultant combustion products are expanded.

As the compressor section draws in atmospheric air and compresses it,the air from the compressor section is suitable for provision to theenvironmental control system (ECS) of the aircraft. In existing ECSconfigurations, air is bled from the compressor section at a temperatureand a pressure in excess of the temperature and pressure required by theECS and is conditioned using a pre-cooler. After being pre-cooled theair is provided to the ECS, and excess pressure is dumped from the ECS.The excess pressure dump results in an overall efficiency loss to theengine.

SUMMARY OF THE INVENTION

In one exemplary embodiment an engine driven environmental controlsystem (ECS) air circuit includes a gas turbine engine including acompressor section, the compressor section including a plurality ofcompressor bleeds, a selection valve selectively connecting each of saidbleeds to an input of an intercooler, and a second valve configured toselectively connect an output of said intercooler to at least oneauxiliary compressor, the output of each of the at least one auxiliarycompressors being connected to an ECS air input.

In another example of the above described engine driven ECS air circuitthe at least one auxiliary compressor comprises a plurality of auxiliarycompressors.

In another example of any of the above described engine driven ECS aircircuits at least one of said compressor bleeds is a compressor bleedpositioned between a low pressure compressor and a high pressurecompressor.

In another example of any of the above described engine driven ECS aircircuits the intercooler is an air to air heat exchanger.

In another example of any of the above described engine driven ECS aircircuits a heat sink of the air to air heat exchanger is fan air.

Another example of any of the above described engine driven ECS aircircuits further includes an aircraft controller controllably connectedto the selection valve and to the second valve such that the aircraftcontroller controls a state of the selection valve and a state of thesecond valve.

In another example of any of the above described engine driven ECS aircircuits the aircraft controller includes a memory storing instructionsconfigured to cause the controller to connect a bleed having a requiredflowrate for an ECS operating requirement, and wherein the connectedbleed has a pressure requirement below a pressure requirement of the ECSinlet.

In another example of any of the above described engine driven ECS aircircuits the at least one auxiliary compressor comprises a plurality ofauxiliary compressors and wherein the aircraft controller includes amemory storing instructions configured to cause the controller toalternate auxiliary compressors operating as a primary compressor on aper flight basis.

In another example of any of the above described engine driven ECS aircircuits the plurality of compressor bleeds comprises at least fourbleeds.

In another example of any of the above described engine driven ECS aircircuits at least one of said at least one auxiliary compressorsincludes an electric motor, and wherein the electric motor is configuredto drive rotation of the corresponding auxiliary compressor.

In another example of any of the above described engine driven ECS aircircuits at least one of said at least one auxiliary compressor includesa mechanical motor, and wherein the mechanical motor is configured todrive rotation of the corresponding auxiliary compressor.

An exemplary method for supplying engine air to an environmental controlsystem (ECS) includes selecting compressor bleed from a plurality ofcompressor bleeds, the selected compressor bleed providing air at ahigher temperature than a required ECS inlet air temperature maximum andat a lower pressure than a required ECS inlet air pressure, cooling thebleed air from the selected bleed using an intercooler such that thebleed air is below the required ECS inlet air temperature maximum,compressing the bleed air using at least one auxiliary compressor suchthat the bleed air is at least the required ECS inlet air pressure, andproviding the cooled compressed bleed air to an ECS air inlet.

In another example of the above described exemplary method for supplyingair to an ECS bleed air is cooled by the intercooler prior to becompressed, thereby decreasing a magnitude of work required to compressthe bleed air to the desired pressure.

In another example of any of the above described exemplary methods forsupplying air to an ECS compressing the bleed air using the at least oneauxiliary compressor comprises driving rotation of the at least oneauxiliary compressor via an electric motor.

In another example of any of the above described exemplary methods forsupplying air to an ECS selecting a compressor bleed from a plurality ofcompressor bleeds comprises selecting a corresponding compressor bleedfrom each of multiple engines simultaneously.

In another example of any of the above described exemplary methods forsupplying air to an ECS compressing the bleed air using at least oneauxiliary compressor comprises simultaneously operating at least twoauxiliary compressors in response to at least one of the enginesshutting down.

In another example of any of the above described exemplary methods forsupplying air to an ECS compressing the bleed air using at least oneauxiliary compressor further comprises alternating a primary compressorbetween a plurality of auxiliary compressors on a per flight basis.

These and other features of the present invention can be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates an exemplary gas turbine engine.

FIG. 2 schematically illustrates an electro-pneumatic environmentalcontrol system (ECS) air circuit for an aircraft.

DETAILED DESCRIPTION OF AN EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10668 meters). The flight condition of 0.8 Mach and35,000 ft (10668 m), with the engine at its best fuel consumption—alsoknown as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—isthe industry standard parameter of lbm of fuel being burned divided bylbf of thrust the engine produces at that minimum point. “Low fanpressure ratio” is the pressure ratio across the fan blade alone,without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressureratio as disclosed herein according to one non-limiting embodiment isless than about 1.45. “Low corrected fan tip speed” is the actual fantip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram ° R)/(518.7 ° R)]{circumflex over ( )}0.5. The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5 m/s).

In order to provide air from the compressor section 24 to the aircraftenvironmental control system (ECS), multiple bleeds are incorporated inthe compressor section 24 (illustrated schematically in FIG. 2). Each ofthe bleeds withdraws air from the compressor section 24 at a givencompressor stage according to known aircraft bleed techniques and usingknown bleed apparatuses. Contemporary aircraft systems for providing airto an ECS bleed air from a stage necessary to meet a required flow rateof the ECS. Bleeding at these stages, however, necessitates bleeding airat a temperature that is in excess of a maximum allowable temperature,and at a pressure that is in excess of a maximum allowable pressure forthe ECS. In order to reduce the temperature, a pre-cooler heat exchangeris positioned in the air circuit and reduces the temperature of thebleed air before the bleed air is provided to the ECS. Once at the ECS,the excess pressure is dumped, resulting in air provided to the ECS thatmeets the temperature, pressure and flow requirements. Pressurization ofthe air passing through the compressor section 24 requires energy, andthe provision of excess pressure to the ECS constitutes waste, anddecreases the efficiency at which the engine 20 can be operated.

FIG. 2 schematically illustrates an electro-pneumatic ECS air circuit100 that reduces the inefficiencies associated with providing air from acompressor to an ECS. The electro-pneumatic ECS air circuit 100 includesmultiple bleeds 102, 104, 106, 108 within a compressor section 122 of anengine 120. Each of the bleeds 102, 104, 106, 108 is connected to anintercooler 130 via a selection valve 140. The intercooler 130 operatesas a heat exchanger to cool the bleed air. In the exemplaryillustration, the bleeds 102, 104, 106, 108 are positioned at aninter-compressor stage between a low pressure compressor 122 a, and ahigh pressure compressor 122 b (bleed 102), and at a high pressurecompressor 122 b third stage (bleed 104), 6^(th) stage (bleed 106), and8^(th) stage (bleed 108). In alternative example engines, the bleedlocations can be positioned at, or between, alternative compressorstages, depending on the specific flow, temperature, and pressurerequirements of the aircraft incorporating the engine 120. In yetfurther alternative example engines 120, alternative numbers of bleedscan be utilized depending on the specific requirements of the aircraft.

An aircraft controller 101 controls the selection valve 140 such that,at any given time, air is provided from a bleed 102, 104, 106, 108having the appropriate flow requirements of the ECS at the currentoperating conditions of the aircraft. While the bleed 102, 104, 106, 108selected by the controller 101 provides air at acceptable flow levels,the bleed 102, 104, 106, 108 is selected to provide air that is underpressured. In other words, the pressure of the air provided by theselected bleed 102, 104, 106, 108 is below the pressure required by theECS. Further, the air selected exceeds the temperature requirements ofthe ECS.

After passing through the selection valve 140, the air is passed to theintercooler 130. The intercooler 130 is a heat exchanger that cools thebleed air prior to providing the air to the ECS. The exemplaryintercooler 130 utilizes fan air, provided from the bypass flowpath ofthe engine 120, to cool the air in a conventional air to air heatexchanger format. In alternative examples, alternative style heatexchangers can be utilized as the intercooler 130 to similar effect.

Cooled air from the intercooler 130 is provided to a second valve 150.The second valve 150 is controlled by the aircraft controller 101 andprovides air to a first auxiliary compressor 160, a second auxiliarycompressor 162, or both the first and second auxiliary compressor 160,162. Each of the auxiliary compressors 160, 162 is driven by acorresponding electric motor 164, 166 and raises the pressure of the airto a required pressure level for provision to the ECS. In alternativeexamples, one or both of the electric motors 164, 166 can be replaced orsupplemented by a mechanical motor and/or a mechanical connection to arotational source within the engine 120 or within the aircraftincorporating the engine 120. Once pressurized via the auxiliarycompressors 160, 162 the air is provided to the ECS. In alternativeexamples, a single auxiliary compressor 160 can be used in place of thefirst and second auxiliary compressors 160, 162. In yet furtheralternative examples, three or more auxiliary compressors can beincluded, with the controller 101 rotating between the auxiliarycompressors as necessary.

By cooling the bleed air prior to providing the bleed air to auxiliarycompressors 160, 162, the amount of work required to compress the air atthe auxiliary compressor 160, 162 is reduced, thereby achieving a fuelefficiency savings.

While the circuit 100 is illustrated in FIG. 2 with a single engine 120,a similar circuit can be utilized with multiple engines 120, with theair from the bleeds 102, 104, 106, 108 of each engine 120, being mixedafter being cooled in a corresponding intercooler 130. Alternatively,the air from each engine 120 can be mixed at alternate positions in theECS air circuit 100 prior to provision to auxiliary compressors 160,162.

In the exemplary circuit 100 only one of the auxiliary compressors 160,162 is required to provide sufficient pressurization to the ECS duringstandard operating conditions. As such, only a single auxiliarycompressor 160, 162 is typically operated during a flight. In order toeven out wear between the auxiliary compressors 160, 162 the primaryoperating auxiliary compressor 160, 162 is alternated between flights ona per flight basis. Alternating between auxiliary compressors 160, 162further allows earlier detection, and correction, of a damaged orinoperable second auxiliary compressor 162.

During flight, when one engine 120 shuts down, either due to mechanicalfailure, or for any other reason, the air provided from the bleeds 102,104, 106, 108, is reduced proportionally. By way of example, if thereare two engines 120, and one shuts down, the air provided to theauxiliary compressors 160, 162 is cut in half. In order to remedy this,in the exemplary system when one engine 120 shuts down, the currentlyinactive auxiliary compressor 160, 162 begins operating simultaneouslywith the currently operating auxiliary compressor 160, 162. Thesimultaneous operations ensure that any lost pressure due to the loss ofan engine is compensated for using air from the operating engine orengines. In aircraft having more than two auxiliary compressors 160,162, the controller 101 can apply a proportional control to one or moreof the auxiliary compressors to ensure that adequate pressure ismaintained at the ECS in proportion to the pressure lost due to the lackof operation of the engine.

It is further understood that any of the above described concepts can beused alone or in combination with any or all of the other abovedescribed concepts. Although an embodiment of this invention has beendisclosed, a worker of ordinary skill in this art would recognize thatcertain modifications would come within the scope of this invention. Forthat reason, the following claims should be studied to determine thetrue scope and content of this invention.

1. A method for supplying engine air to an environmental control system (ECS) comprising: Selecting a compressor bleed from a plurality of compressor bleeds using a selection valve connected to each compressor bleed in the plurality of compressor bleeds, the plurality of compressor bleeds including a first bleed positioned at an intercompressor stage between a low pressure compressor and a high pressure compressor, a second bleed positioned at a third stage of the high pressure compressor, a third bleed positioned at a sixth stage of the high pressure compressor, and a fourth bleed positioned at an eighth stage of the high pressure compressor, the selected compressor bleed providing air at a higher temperature than a required ECS inlet air temperature maximum and at a lower pressure than a required ECS inlet air pressure; connecting the selected compressor bleed to an input of an intercooler using the selection valve; cooling the bleed air from the selected bleed using the intercooler such that the bleed air is below the required ECS inlet air temperature maximum; connecting an output of the intercooler to at least one auxiliary compressor using a second valve, wherein the output of every auxiliary compressor in the at least one auxiliary compressor is connected to an ECS air inlet; compressing the bleed air using the at least one auxiliary compressor such that the bleed air is at least the required ECS inlet air pressure; and providing the cooled compressed bleed air to the ECS air inlet.
 2. The method of claim 1, wherein compressing the bleed air using the at least one auxiliary compressor comprises driving rotation of the at least one auxiliary compressor via an electric motor.
 3. The method of claim 1, wherein selecting a compressor bleed from a plurality of compressor bleeds comprises selecting a corresponding compressor bleed from each of multiple engines simultaneously.
 4. The method of claim 3, wherein compressing the bleed air using at least one auxiliary compressor comprises simultaneously operating at least two auxiliary compressors in response to at least one of the engines shutting down.
 5. The method of claim 1, wherein compressing the bleed air using at least one auxiliary compressor further comprises alternating a primary compressor between a plurality of auxiliary compressors on a per flight basis.
 6. A method for supplying engine air to an environmental control system (ECS) comprising: selecting compressor bleed from a plurality of compressor bleeds, the selected compressor bleed providing air at a higher temperature than a required ECS inlet air temperature maximum and at a lower pressure than a required ECS inlet air pressure; cooling the bleed air from the selected bleed using an intercooler such that the bleed air is below the required ECS inlet air temperature maximum; compressing the bleed air using at least one auxiliary compressor such that the bleed air is at least the required ECS inlet air pressure; and providing the cooled compressed bleed air to an ECS air inlet.
 7. The method of claim 6, wherein bleed air is cooled by the intercooler prior to be compressed, thereby decreasing a magnitude of work required to compress the bleed air to the desired pressure.
 8. The method of claim 6, wherein compressing the bleed air using the at least one auxiliary compressor comprises driving rotation of the at least one auxiliary compressor via an electric motor.
 9. The method of claim 6, wherein selecting a compressor bleed from a plurality of compressor bleeds comprises selecting a corresponding compressor bleed from each of multiple engines simultaneously.
 10. The method of claim 9, wherein compressing the bleed air using at least one auxiliary compressor comprises simultaneously operating at least two auxiliary compressors in response to at least one of the engines shutting down.
 11. The method of claim 6, wherein compressing the bleed air using at least one auxiliary compressor further comprises alternating a primary compressor between a plurality of auxiliary compressors on a per flight basis.
 12. The method of claim 6, wherein the plurality of compressor bleeds includes a first bleed positioned at an intercompressor stage between the low pressure compressor and the high pressure compressor, a second bleed positioned at a third stage of the high pressure compressor, a third bleed positioned at a sixth stage of the high pressure compressor, and a fourth bleed positioned at an eighth stage of the high pressure compressor.
 13. The method of claim 6, wherein the plurality of compressor bleeds includes a bleed positioned at an intercompressor stage between the low pressure compressor and the high pressure compressor.
 14. The method of claim 6, wherein the plurality of compressor bleeds includes a bleed positioned at a third stage of a high pressure compressor.
 15. The method of claim 6, wherein the plurality of compressor bleeds includes a bleed positioned at a sixth stage of a high pressure compressor.
 16. The method of claim 6, wherein the plurality of compressor bleeds includes a bleed positioned at an eighth stage of a high pressure compressor. 